Propellant transfer system and method for resupply of fluid propellant to on-orbit spacecraft

ABSTRACT

Herein is disclosed a propellant transfer system and method for refueling on-orbit spacecraft. The system and method are configured to allow for resupply of spacecraft configured to be fueled by either a bipropellant (oxidizer and fuel) or a monopropellant (typically hydrazine). The system and method are particularly suited for resupply of satellites not originally prepared for refueling as well but the system may also be used for as satellites specifically designed for refueling.

CROSS REFERENCE TO RELATED U.S. PATENT APPLICATIONS

This patent application is a divisional application of U.S. patentapplication Ser. No. 13/678,281 claiming the priority benefit from U.S.provisional patent application Ser. No. 61/559,801 filed on Nov. 15,2011, entitled PROPELLANT TRANSFER SYSTEM AND METHOD FOR RESUPPLY OFFLUID PROPELLANT TO ON-ORBIT SPACECRAFT, filed in English, all of whichare incorporated herein in their entirety by reference.

FIELD

The present disclosure relates to a method and system for on-orbit fluid(i.e., liquid or gas) propellant resupply (or propellant removal) ofartificial satellites either near the end of their originally scheduledlifetime or at any point in their life after an anomalous orbitinsertion, another commissioning problem, or greater than plannedmaneuvering during nominal operations. More particularly the system andmethod are designed for resupply of satellites not originally preparedfor refueling as well as satellites designed for refueling.

BACKGROUND

Many satellites currently in operation were designed with a finiteamount of propellant and were not designed for the possibility of beingresupplied with propellant. The design philosophy relied uponreplacement of the satellites after they had exhausted the on-boardpropellant supply. In view of the expense of replacing satellites, itwould be very advantageous to be able to resupply satellites withpropellant which are either near their end of propellant life butotherwise functional, or have suffered an infant propulsion systemfailure or insertion anomaly, or have been maneuvered more thanoriginally intended for their nominal operations, thereby extendingtheir operational life by several or many years.

It is estimated that as many as half of all GEO communication satellitesend their 10 to 15 year life with all or most of their subsystems stillfunctional and it is only the depletion of the carefully budgetedpropellant load that drives retirement of the satellite. Using a currenteconomic model, the ability to refuel several of these end-of-lifesatellites in a single mission would cost-effectively extend each oftheir useful lives by 3 to 5 years and thereby delay the need to outlaythe very high capital costs to launch a replacement for each satelliteif desired. Some satellites suffer from primary propulsion systemfailures or launch vehicle upper stage related failures soon after theyare launched. In these cases the entire book value must be written offand compensation paid to the operator by the insurer. The satellitebecomes an asset of the insurer and will eventually have to be disposedof in a graveyard or re-entry orbit. If these assets can be resuppliedwith propellant, enabling them to transfer to an orbital station ingeosynchronous orbit and extending their life by 5 to 10 years, most orall of the value of the satellite can be recovered.

The key technical difficulty is that these satellites were not designedfor robotic servicing, and it is not generally accepted that suchmissions are technically possible. Specifically, most satellites aredesigned with propellant fill and drain valves that were intended to befilled once prior to launch and never opened or manipulated followinglaunch. Thus, accessing these fill and drain valves remotely in-orbitpresents several major challenges and would involve several operations,each of which is difficult to accomplish robotically including: cuttingand removal of the protective thermal blankets, removal of severallockwires hand wrapped around the valves, unthreading and removing outerand inner valve caps, mating a propellant fill line to the valve nipple,mechanically actuating the valve and, when propellant resupply iscomplete, replacing the inner valve cap.

On-orbit servicing has been the subject of much study over the pastthirty years. The idea of maintaining space assets, rather thandisposing of and replacing them, has attracted a variety of ideas andprograms. So far, the concept has only found a home in the manned spaceprogram where some success can be attributed to the Solar Max and HubbleSpace Telescope repair missions, Palapa-B2 and Westar rescue missions,and the assembly and maintenance of the International Space Station.

Until recently there have been no technologies disclosed that can solvethe problem of accessing the propellant system of an unpreparedsatellite for the purpose of replenishing station-keeping propellant.The majority of satellites in orbit today were not designed with orbitalpropellant resupply in mind and access to the propellant system isdesigned to be accessed by a human on earth before launch. Thetechnologies required to access the client spacecraft's propellantsystem for the purposes of propellant resupply still have a very lowtechnology readiness level, and are generally considered to be the mainobstacle to a successful servicing mission.

Transferring propellants used for spacecraft propulsion systems from onesource to another can be very dangerous due to the corrosive andexplosive nature of many of the fluids involved. For example,inadvertent mixing of fuel and oxidizer in bipropellant systems willcause immediate combustion, so a fluid transfer system for bipropellantneeds to ensure that no accidental mixing occurs.

Therefore, it would be very advantageous to provide a propellanttransfer system for transferring propellant from a servicing spacecraftto a client satellite which has flexibility to deliver propellant usingmore than one modality depending on the circumstances of the satellite,propellant system parameters, and the like. It would be veryadvantageous for such a system to able to able to transfer bipropellantsin addition to monopropellants, pressurants, and ion or plasmapropulsion propellants.

SUMMARY

The present disclosure relates to a propellant transfer system andmethod for on-orbit propellant resupply of an artificial satellite. Thesystem and method are configured to allow for propellant resupply of asatellite configured to use either a bipropellant oxidizer and fueland/or a monopropellant and/or a pressurized gas propellant orpressurant. The system and method are particularly suited for propellantresupply of satellites not originally prepared for propellant resupplybut the system may also be used for satellites specifically designed forpropellant resupply later in their operational life.

Disclosed is a system mounted on a servicer spacecraft for transferringfluid to a client satellite, the client satellite including at least onestorage tank having at least a first fluid transfer coupling,comprising:

a) a fluid storage and routing system for storing and routing fluid fromthe servicing spacecraft to the at least one storage tank on the clientsatellite, the storage and routing system including

-   -   at least one pressurized tank containing a pressurized gas, at        least one fluid storage tank, and an associated fluid transfer        coupling,    -   first and second flow paths connecting the at least one fluid        storage tank, the at least one pressurized gas tank and the        associated fluid transfer coupling; and

b) a flow control system configured for

-   -   detecting and adjusting pressure and flow rate of the gas and        fluid through the first and second flow paths,    -   detecting pressure in the at least one storage tank on the        client satellite once the associated fluid transfer coupling is        coupled to the fluid transfer coupling on the client satellite,

transferring fluid from said fluid tank, using pressurized gas, to saidat least one storage tank on the client satellite via said first flowpath at a pressure less than or equal to a pressure of the at least onepressurized gas tank, or via said second flow path at a pressure greaterthan a pressure of the at least one pressurized gas tank.

There is also disclosed system mounted on a servicer spacecraft fortransferring fluid to a client satellite, the client satellite includingat least one storage tank having at least a first fluid transfercoupling, comprising:

a) means for storing and routing fluid from said servicing spacecraft tothe at least one storage tank on the client satellite; and

b) means for controlling flow of gas and fluid, said means forcontrolling flow of gas and fluid configured for

-   -   detecting and adjusting pressure and flow rate of said gas and        fluid through said first and second flow paths,    -   detecting pressure in the at least one storage tank on the        client satellite once said associated fluid transfer coupling is        coupled to the fluid transfer coupling on the client satellite,    -   transferring fluid from said fluid tank to said least one        storage tank on the client satellite using the pressurized gas        -   via said first flow path at a pressure less than or equal to            a pressure of the at least one pressurized gas tank, or        -   via said second flow path at a pressure greater than a            pressure of the at least one pressurized gas tank.

There is also disclosed a propellant transfer system mounted on aservicer satellite for transferring bipropellant and/or monopropellantbetween the servicer satellite and propellant tank of client satellitesconfigured for propulsion with bipropellant and/or monopropellantrespectively, each client satellite having a fill/drain valve associatedwith each propellant tank for accessing the propellant storage tank,comprising:

a) first, second and third propellant transfer subsystems, eachpropellant transfer subsystem including,

-   -   at least one pressurized tank containing a pressurized gas, at        least one propellant storage tank, the at least one propellant        storage tank and the at least one pressurized gas tank being in        flow communication with each other through a routing tube        system;

b) a flow control system integrated with the interconnecting tube systemfor

-   -   detecting and adjusting pressure and flow rate of the gas and        propellant, and    -   detecting pressure in the client satellite propellant storage        tank once the propellant transfer subsystem is coupled to the        client satellite and the fill/drain valve on the client        satellite propellant tank is open,

c) a command and control system interfaced with the flow control systemof each propellant transfer subsystem, the command and control systembeing configured to regulate pressure and flow rate of propellantbetween each propellant transfer subsystem and associated clientsatellite propellant tanks based on a detected pressure of the clientsatellite propellant tank; and

d) a communication system configured to provide communication betweenthe command and control system and a remote operator for remoteteleoperator control, or a mixture of teleoperator control andsupervised autonomy control, or fully autonomous control of propellanttransfer operations between the first, second and third propellanttransfer subsystems and the client satellite propellant tanks and theassociated propellant tanks on the client satellite.

Also disclosed is a method of refueling a client satellite, comprising:

a) maneuvering a servicer spacecraft into close proximity to a clientsatellite in a position and orientation suitable to releasibly couplethe client satellite to the servicer spacecraft, releasibly coupling theclient satellite to the servicer satellite;

b) deploying, and commanding, a robotic arm mounted on the servicerspacecraft to releasibly grasp a multifunction tool, commanding therobotic arm to manipulate the multifunction tool to access a fill/drainvalve on the client satellite in flow communication with a storage tankon the client satellite;

c) commanding the robotic arm to sequester the multifunction tool andreleasibly grasp a refueling tool, commanding the robotic arm tomanipulate the refueling tool to releasibly grasp a propellant outlethose connected to a propellant transfer system and mate it to thefill/drain valve;

e) opening the fill/drain valve and measuring a pressure in the clientsatellite storage tank and based on the measured pressure, configuring aflow control system mounted on the servicer satellite to dispensepropellant under regulated pressure and flow rate conditions suitablefor the measured pressure;

f) commanding the configured flow control system of the propellanttransfer system to transfer propellant from a propellant storage tank onthe servicer spacecraft through a piping system to the propellant hoseto the storage tank on the client satellite;

g) once a desired quantity of propellant has been transferred to theclient satellite, commanding the robotic arm to manipulate the refuelingtool to close the fill/drain valve;

h) demate the propellant hose from the fill/drain valve and sequesterit; and

i) sequestering the robotic arm and decoupling the servicer spacecraftfrom the client satellite.

A further understanding of the functional and advantageous aspects ofthe disclosure can be realized by reference to the following detaileddescription and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described, by way of example only, withreference to the drawings, in which:

FIG. 1 shows a simple block diagram indicating how the entire propellanttransfer system shown in FIGS. 1 a to 1 f relate to each other.

FIG. 2 shows the schematic drawing of FIG. 1 assembled in one Figure butnow identifying various optional subsystems.

FIGS. 2 a to 2 h show expanded views of sections of the schematic ofFIG. 2;

FIG. 3 shows a schematic drawing of propellant transfer system but nowbut absent all optional and redundant subsystems.

FIG. 4 is an enlarged view of the high pressure pressurant subsystemforming part of the present system.

FIG. 5 shows a part of the block diagram of FIG. 1 but showing only thepropellant transfer system for resupplying client satellites configuredto use a bipropellant, including all redundant and optional subsystems.

FIGS. 5 a to 5 d show expanded views of sections of the schematic ofFIG. 5;

FIG. 6 shows a schematic drawing of a portion of the propellant transfersystem configured for a first method of the propellant transfer processin which propellant is transferred directly by pressure gradient from ahigher pressure propellant storage tank on the servicer spacecraft to alower pressure client satellite propellant tank.

FIGS. 7 a and 7 b shows schematic drawing of a portion of the systemconfigured for a second method of the propellant transfer process inwhich propellant is first cycled from the propellant storage tank to thetransfer tank (FIG. 7 a) and then from the transfer tank to the clientsatellite propellant tank (FIG. 7 b).

FIG. 8 shows the refueling system mounted on a servicer spacecraft witha computer control system in communication with a remote teleoperationcontrol center.

FIG. 9 shows a drain/fill valve located on a client satellite to beresupplied with propellant.

FIG. 10 a shows the backup fill/drain valve with the secondary sealthreaded feature, the backup fill/drain valve actuation nut, the backupfill/drain valve torque reaction feature and the backup fill/drain valveseal fitting.

FIG. 10 b is an elevation view of a replacement secondary seal fittingused to close off the fill/drain valve of the propulsion system on theclient satellite once propellant transfer operations have beencompleted.

FIG. 11 shows the backup fill/drain valve engaged with the clientsatellite drain/fill valve on the client satellite being resupplied.

FIG. 12 shows an exemplary, non-limiting computer control system formingpart of the system disclosed herein.

FIGS. 13 and 13 a shows an alternative embodiment of a fluid transfersystem using a mechanical pump.

DETAILED DESCRIPTION

Various embodiments and aspects of the disclosure will be described withreference to details discussed below. The following description anddrawings are illustrative of the disclosure and are not to be construedas limiting the disclosure. Numerous specific details are described toprovide a thorough understanding of various embodiments of the presentdisclosure. However, in certain instances, well-known or conventionaldetails are not described in order to provide a concise discussion ofembodiments of the present disclosure.

As used herein, the terms, “comprises” and “comprising” are to beconstrued as being inclusive and open ended, and not exclusive.Specifically, when used in this specification including claims, theterms, “comprises” and “comprising” and variations thereof mean thespecified features, steps, or components are included. These terms arenot to be interpreted to exclude the presence of other features, steps,or components.

As used herein, the term “exemplary” means “serving as an example,instance, or illustration,” and should not be construed as preferred oradvantageous over other configurations disclosed-herein.

As used herein, the terms “about” and “approximately”, when used inconjunction with ranges of dimensions of particles, compositions ofmixtures, or other physical properties or characteristics, are meant tocover slight variations that may exist in the upper and lower limits ofthe ranges of dimensions so as to not exclude embodiments where onaverage most of the dimensions are satisfied but where statisticallydimensions may exist outside this region. It is not the intention toexclude embodiments such as these from the present disclosure.

As used herein the term or “satellite” being resupplied with propellantrefers to artificial satellites.

As used herein, the phrase “backup fill/drain valve” refers to valveattached to the client satellites fill/drain valve that can be closedand left behind should it prove not possible to properly close theclient fill/drain valve (i.e., a leak through the client fill/drainvalve is detected) after the completion of propellant transferoperations.

Definitions of Component Symbols

The following list of parts and their function is given below.

CV refers to Check Valves and are used to prevent backflow of gas (whichmay be contaminated with propellant vapor) back to the tubing downstreamof the regulators and upstream of the check valves. Mixing of oxidizervapor with fuel or monopropellant vapor must be prevented as the twowould react on contact.

F—refers to Flow Meter transducers and are used to measure thepropellant flow rate. In the prototype an ultrasonic flow meter isemployed, which achieves the desired flow rate measurement accuracy andavoids any contact of sensitive transducer parts to corrosivepropellants.

FDV—refers to Fill/Drain Valves and are used to load/unload fluid(propellant) into/from the various propellant tanks.

FVV—refers to Fill/Vent Valves and are used to load/unload gas(pressurant) into/from the pressurant tanks, or are used to access thetubing for pressure testing.

FG—refers to gas filters which are used to prevent particulatecontaminates in the pressurant tanks or form outside the system frommigrating downstream where it might otherwise interfere with valveoperation.

FL—refers to Filters and are used to prevent particulate contaminates inthe storage tanks from migrating downstream where it might otherwiseinterfere with valve operation.

HPLV—refers to High Pressure Latch Valve and are used to isolate thedownstream regulators from the high pressure source during launch.

HPT—refers to High Pressure Transducer and are used to measure thepressure of the high pressure portion of the system.

IVG—refers to Gas Isolation Valves and are used to control the flow ofgas from the high pressure section to the transfer tank (IVG₁), betweenthe storage tank and the transfer tank (IVG₂), and from the transfertank to space (IVG₃).

IVL—refers to Liquid Isolation Valve and are used to control the flow offluid (propellant) between the storage tank and the transfer tank/(or apump) (IVG₁), and between the transfer tank/pump of the propellanttransfer system on the servicer spacecraft and the propellant tank onthe client satellite (IVG₂).

LPT—refers to Low Pressure Transducer and are used to measure thepressure at various parts of the system.

NTD—refers to non-thrusting devices and are used to vent pressurant tospace without generating thrust which would otherwise disturb theservicer+client spacecraft attitude and/or orbit.

PR—refers to pressure regulators which act to reduce the upstream gaspressure from a high value (up to about 320 bar) in the pressurantstorage tanks to a lower downstream value (no higher than 30 bar) forpressurizing the propellant storage tanks and operating the transfertanks.

ST—Refers to the propellant Storage Tanks and are used to store all thepropellant to be transferred to the client satellite and potentiallyalso to be used by the servicer spacecraft propulsion system for orbitmaneuvers. They are filled with propellant prior to launch, but can alsobe resupplied with propellant when in orbit.

TT—refers to Transfer Tanks and are used for several purposes includingrelief of the flex hose pressure between client satellite propellanttransfer operations, as a ‘pump’ to transfer propellant cyclically fromthe storage tanks against a pressure gradient to the client, as acalibrated volume to check the flow meter performance, and as a back-upmeans of propellant metering using a simple count of cycles. Thetransfer tanks are small and have separate gas and fluid volumesseparated by a diaphragm or a sealed sliding piston. The transfer tank,in combination with its associated valves, is essentially apneumatically powered positive displacement pump. It could be replacedby some other form of pump such as a motor-driven piston pump, gearpump, or diaphragm pump. These alternative embodiments may be employedparticularly if gas-phase propellants are being used instead of liquidpropellants.

UFM Electronics—refers to the (ultrasonic) flow meter electronics whichare used to interface to the flow meter transducer.

The symbol ▪—refers to thermistors which are used measure thetemperature of associated equipment of the propellant transfer system.

He—refers to gas pressurant which is used to displace fluid propellant.Other non-reactive gases such as nitrogen (N) could also be used, buthelium (He) is preferred for space application due to its lower mass.The Helium is stored at high pressure in pressurant tanks in the highpressure section.

F—refers to liquid bipropellant fuel and is delivered to the clientsatellite. It reacts with oxidizer to produce thrust. A typical liquidfuel is monomethyl hydrazine (MMH), though other fuels such as hydrazine(N₂H₄) or unsymmetric dimethyl hydrazine (UDMH) may also be used.

U—refers to a second chemically compatible liquid bipropellant fuel. Ifthe primary fuel is MMH, then a suitable second fuel would be UDMH.

H—refers to liquid monopropellant and is delivered to clients. Itdisassociates in the presence of a catalyst to produce thrust. It alsocan act like a fuel in that it reacts with oxidizer to produce thrust. Atypical monopropellant is hydrazine (N₂H₄) though other monopropellantssuch as ammonia (NH₃) may also be used.

X—refers to liquid bipropellant oxidizer and is delivered to clients. Itreacts with fuel to produce thrust. A typical liquid oxidizer is amixture of oxides of nitrogen with 97% N₂O₄ and 3% NO (MON3), thoughother oxidizers such as fuming nitric acid may also be used.

FIG. 1 shows a schematic drawing of the complete propellant transfersystem shown generally 10, due to the complexity of the system it isbroken up in sections shown from FIG. 1 a to 1 f. The component labelsassociated with all valves, sensors, pressure transducers, etc. aredefined above. All redundant components, such as valves, pressuretransducers and the like shown in FIGS. 1 to 7 b have a ‘B’ in thesubscript of their labels (as compared to ‘A’ for the equivalent primecomponents). In addition to these Figures showing the redundantcomponents, FIGS. 2 a to 2 h, the optional subsystems of system 10 areshown.

FIG. 1 shows a schematic drawing of the complete propellant transfersystem shown generally at 10. The components associated with all valves,sensors, pressure transducers, etc. are defined above. All redundantcomponents, such as valves, pressure transducers and the like shown inFIGS. 1 to 7 b have a ‘B’ in the subscript of their labels (as comparedto ‘A’ for the equivalent prime components). In addition to theseFigures showing the redundant components, FIG. 2, the optionalsubsystems of system 10.

For example, the propulsion interfaces 120 associated with thebipropellant transfer subsystems, and propulsion interface 122associated with the monopropellant transfer subsystem are optional.These interfaces supply bipropellant (fuel and oxidizer) and/ormonopropellant to the servicer spacecraft propulsion system. Thisreduces the overall mass of the servicer spacecraft by avoiding the needfor separate propellant tanks in the propulsion system. It also allowsthe propellant to either be delivered to client spacecraft or to be usedby the servicer spacecraft for orbit maneuvers. Such orbit maneuvers maybe used, for example, to rendezvous with the client spacecraft or evento change the client spacecraft orbit (while the servicer spacecraft isdocked to the client spacecraft).

Further, the cold gas thruster subsystem 102 is also optional. Cold gasthrusters could be useful to arrest and/or reverse the final approach ofthe servicer spacecraft to the client spacecraft just prior to dockingin the event of an anomaly. Using the high pressure supply of thepropellant transfer system would avoid the need for a second source ofgas for cold gas thrusters.

The alternative propellant string 106 for the fuel (and/or oxidizer)transfer subsystem is also optional. One or more of them could provide ameans of delivering additional chemically-compatible fuel (or oxidizer)types without the addition of other complete propellant transfersubsystems 14 (or 12). For the example, using MMH (monomethyl hydrazine)as the primary fuel, a potential secondary fuel would be UDMH(unsymmetric dimethyl hydrazine), but other secondary fuel types arealso possible depending on the primary fuel. Small amounts of these twofuels would mix in the transfer tank/pump and downstream components whenalternating between delivering the two, but this mixing is typicallytolerable within the fuel purity specifications defined by the militarystandards for these bipropellant fluids.

The Helium ORU (Orbital Replaceable Unit) interface 110 is optional. Itprovides a fluid interface through which additional pressurant could beprovided to the servicer spacecraft thereby extending its useful life.Fluid propellant can be resupplied to the propellant transfer system byventing pressurant from the propellant storage tanks and then reversingthe propellant transfer process to transfer propellant from some otherspacecraft to the servicer spacecraft. The other spacecraft can be aretired one that still has propellant remaining or it can be apurpose-built ‘tanker’ spacecraft. Resupplying pressurant is morechallenging due to the high pressure in the servicer spacecraftpressurant storage tanks. It is proposed, therefore, that pressurant beprovided at low pressure through the He ORU interface. The He ORU wouldbe carried to the servicer spacecraft by the tanker spacecraft and thentransferred to the servicer using the servicer's robotic arm.

The pressurant resupply lines 100 are optional. These lines and theassociated redundant valves (IVG₄ and IVG₅) would be used to resupplythe client with pressurant should the pressure in the client propellanttanks be too low. For clients with PMD (propellant management device)type propellant tanks, pressurant resupply is possible through theclient satellite propellant fill/drain valve as there is no physicalbarrier between the propellant and pressurant. The pressurant bubble inthe propellant line would be flushed out by propellant during subsequentpropellant resupply. For clients with bladder or diaphragm type tanks,pressurant resupply would have to be through an appropriate pressurantfill/vent valve but only if the pressure here is within the capabilityof the propellant transfer system.

FIG. 3 shows a schematic drawing of the propellant transfer system 10absent the optional subsystems 100, 102, 106, 110, 120 and 122 discussedabove. The propellant transfer system 10 includes a first propellanttransfer subsystem 14 for resupplying the fuel tank of satellitesconfigured to use a bipropellant (fuel and oxidizer) system andcontaining a second subsystem 12 for resupplying the oxidizer tank ofthe client satellite. Propellant transfer system 10 includes a thirdpropellant transfer subsystem 16 configured to refuel a storage tank 36on a satellite configured to use a monopropellant. FIG. 3 also shows allthe subsystems 12, 14 and 16 coupled to the respective propellant tankon the satellite being refueled, specifically oxidizer transfersubsystem 12 coupled to propellant tank 22 for storing the oxidizer andsubsystem 14 coupled to a propellant tank 24 for storing the fuel on asatellite configured to use the bipropellant. Subsystem 16 is showncoupled to a propellant tank 26 for storing monopropellant on asatellite configured to use a monopropellant. In all cases, whethertransferring monopropellant or the bipropellant, a backup fill/drainvalve 52 is first mated to the drain/fill valves of the clientsatellites. Backup fill/drain valve 52 will be discussed in more detailhereafter.

Subsystems 12 and 14 each contains two (2) flow meter transducers andstorage tanks 30 a, 30 b are identical as can be seen from FIG. 3. Eachof subsystems 12 and 14 are configured with redundancy options as notedabove so each contains two (2) sets of tubing, valves, and pressuretransducers, and transfer tanks 40 a and 40 b. Pressurant gas flows fromthe high pressure helium storage tank 44 to the storage tanks 30 a, 30 band hence to the outlet hoses 32 and 34. As previously mentioned, thetransfer tanks, in combination with their associated valves, areessentially pneumatically powered positive displacement pump mechanisms.

Subsystem 16 configured for holding and transferring monopropellant hasessentially the same valving/tubing arrangement as for each of thesubsystems 12 and 14. Each subsystem 12, 14 and 16 includes its own flowcontrol system comprised of various valves, leak detectors, gas pressureregulators, pressure transducers, and flow sensors and meters stationedin strategic locations in the routing tube system interconnecting thehigh pressure tank 44 to the propellant storage tanks 30 a, 30 b, 36,the transfer tanks 40 a, 40 b, and 42 and the outlet hoses 32, 36. Thesymbols for all these components of which the flow control system iscomprised are defined above.

Thus, various storage tanks and routing tube system and pump mechanism,shown in the Figures provide a means for storing and routing fluid fromthe servicing spacecraft to associated storage tanks on the clientsatellite. The plurality of valves, leak detectors, pressure sensors,gas pressure regulators, temperature sensors flow sensors and metersinterfaced with the computer control system and stationed in strategiclocations in the routing tube system provide a means for controlling theflow of fluid (gas and/or liquids) during the transfer process.

The elements of each flow control system are interfaced with a commandand control system for regulating or controlling parameters associatedwith fluid flow such as pressure and flow rates. The command and controlsystem is described in more detail below and may includecomputer/processors mounted on the propellant transfer system. Forsaving mass preferably only one redundant computer is used to controlall the subsystems but it will be understood that each subsystem couldhave its own computer controller.

As mentioned above, subsystems 12 and 14 each include two transfer tanks40 a (prime) and 40 b (redundant) in flow communication with a pair ofstorage tanks 30 a and 30 b in any combination while subsystem 16includes two transfer tanks 42 a and 42 b in flow communication with asingle storage tank 36. This is solely to indicate that either a singleor multiple storage tanks can be used as dictated by the servicerspacecraft configuration. Two or more storage tanks should be used ifthey are mounted off-axis so that they balance each other throughout theservicer spacecraft's lifetime whereas only one storage tank isnecessary if it is on-axis. The storage tanks 30 a and 30 b are in flowcommunication with the high pressure section.

FIG. 4 shows an enlarged view of the high pressure (up to about 300 bar)section of the propellant transfer subsystem of FIG. 1. Helium is storedin one or more pressurant tanks 44 (four are shown in FIGS. 1 to 5, withone clearly shown and three (3) partially hidden). Redundant HPLVs areused to isolate the downstream pressure regulators (PR) from the highpressure for launch. Quad redundant pressure regulators reduce thepressure and supply the rest of the refueling system with helium atlower pressure (about 20 bar).

FIGS. 5 to 5 d show the full propellant transfer system with optionalsubsystems (100, 106, 110, and 120 from FIGS. 2 to 2 h) for thesubsystems 12 and 14 for resupplying satellites with the bipropellant.

In all the propellant transfer subsystems, the gas flow rates areregulated by opening/closing on of the gas isolation valves (IVGs) basedon pressure reading from one or more of the high pressure transducers(HPT) used to measure the pressure of the high pressure portion of thesystem and the low pressure transducers (LPT) used to measure thepressure at various parts of the in the propellant transfer subsystemsseparate from the high pressure section.

Propellant flow rate are regulated indirectly by controlling the gaspressure. However in some circumstances, in order to avoid wastingpressurant by venting it, it may not be preferable to decrease the flowrate this way. Thus, in some circumstances the flow rate may beregulated using only of the existing fluid valves. One way to controlthe propellant flow rate is to use the backup fill/drain valve 52 toregulate the flow rate. The backup fill/drain valve 52 opens partiallydepending on how far it actuator nut 132 (see FIG. 10) is rotated. Thusthe flow rate in this preferred mode is controlled by adjustment ofactuator nut 132, turning it slowly open until the flow rate reached thedesired rate. Preferably backup fill/drain valve 52 has an adjustableorifice and is calibrated such that the command and control system cancarefully adjust the orifice size depending on the calibration and aknowledge of the number of turns of actuator nut 132 performed by arefueling tool (discussed hereinafter) mated to the backup fill/drainvalve 52.

FIG. 6 shows an enlarged view of the monopropellant section of thepropellant transfer subsystem of FIG. 1, and the dotted lines 70 frompressurant tank 44 (not shown in FIG. 6) down through storage tank 36and down to exit hose 34 represents the propellant path for thedelivered propellant using a first embodiment of a method of propellanttransfer from the storage tank 36 to the client tank 26. In this firstmode, the “direct transfer mode”, the propellant transfer system isconfigured to provide regulated transfer of the propellant directly fromstorage tank 36 due to a simple pressure gradient from the storage tankwhich is indicated by the broken line 70 from the storage tank 36 to theoutlet which is connected via the flow meter transducer, IVL₁, IVL₂, theflex hose, the backup fill/drain valve, and the refueling tool to theassociated propellant line of the client satellite (the monopropellantline on the client satellite if it is the monopropellant that is beingtransferred, the fuel line if it is the fuel being transferred, or theoxidizer line if it is the oxidizer being transferred). This is thepreferred mode of operation where the client satellite propellant tank26 is at a pressure below that of the servicer spacecraft's propellantstorage tank 36. The pressure of client satellite storage tank 26(between about 5 bar and 20 bar) depends on the client satellite'sthrusters operating pressure range and its state of life. The maximumstorage tank operating pressure (up to about 30 bar) depends on the tankdesign. In this embodiment the “pump mechanism” is the pressurizedstorage tank 44 itself and associated valves etc.

As fluid propellant flows from the servicer's storage tank 36,pressurant gas He in the storage tank expands into the volume vacated bythe propellant, thereby reducing the storage tank 36 pressure. Thestorage tank 36 can be pressurized, or its pressure maintained, byallowing pressurant gas to flow as indicated by the broken arrow 70 fromthe high pressure tank 44 to the storage tank via the valves IVG1 andIVG2.

As propellant flows into the client propellant tank 26, the propellantdisplaces the gas inside the tank 26 thereby compressing it. Thisreverses the reduction of pressure in the client propellant tank 26 overits life due to pressurant gas expansion as propellant is expended fororbit maintenance. In this way, the conditions of the client satellitereturn to an earlier state both in terms of the quantity of propellantand pressure in its tank 26.

FIGS. 7 a and 7 b show a second mode of operation such that the methodof transferring propellant from storage tank 36 to client storage tank26 using a “pumped transfer via transfer tank/pump” mode. In this modeof operation the propellant transfer system is configured to firsttransfer the propellant using the first half of a (pumping) cycledtransfer of propellant from the storage tank 36 to the transfer tank 42a. Referring to FIG. 7 a, the propellant transfer system fills thetransfer tank 42 a with a desired amount of propellant from storage tank36 as shown by the broken arrow 76. The propellant is pushed by apressure gradient induced by venting pressurant from the gas volume 46(separated and isolated from the fluid volume 48 by diaphragm or piston49) of the transfer tank 42 a to space as shown by the broken arrow 78.

Referring to FIG. 7 b, the propellant is then transferred to the clienttank 26 using a pressure gradient induced by pressurizing the transfertank 40 a from the high pressure He tank 44 (not shown) along the pipingpath indicated by arrow 82 such that the fluid in transfer tank 42 a isdriven into storage tank 26 on the dent satellite as shown by the brokenarrow 80. This is the only possible mode of operation where the clientsatellite propellant tank 26 is at a pressure greater than that of theservicer propellant storage tank up to the set point for the pressureregulators in the high pressure section.

Referring to FIG. 3, it is noted that the combination of the transfer(s)tank 40 a (40 b) or 42 a (42 b) and the five valves IVG1, IVG2, IVG3,IVL1, and IVL2 act as a pump when operated in a coordinated way. Whenoperated in other (differently coordinated) ways these componentsperform other functions. Examples of these other functions include thecapability to relieve the pressure in the hoses 32 and 34 to facilitatedisconnection of the refueling tool 50 from the client satellitefill/drain valve 402 and metering of propellant by counting transfercycles. Such metering can be used to verify performance of the flowmeter with or without actual transfer of propellant to a clientsatellite.

It will be appreciated that while the propellant transfer system isprimarily configured for resupply of propellant to a client satellite,it may be used for the reverse process. For example it may be used toresupply of propellants to the servicer spacecraft by venting thestorage tank(s) 36, and then letting the (reverse) pressure gradientpush the propellant from a purpose built tanker spacecraft or from theclient satellite storage tank 26 to the servicer's propellant storagetanks 36. Such an operation would be conducted if the client satelliteis no longer viable for its intended purpose but remains a useful sourceof propellant. This applies equally to both the direct transfer method(FIG. 6) and the cycled transfer method (FIGS. 7 a and 7 b).

For either method of transfer the propellant flow rate is measured usingone or more flow meters integrated into the tube system. Any type offlow meter may be used, for example, but not limited to, ultrasonic flowmeters. The flow rate may be numerically integrated over time by anonboard computer to determine the quantity of propellant transferred.For the ‘pumped’ transfer via the transfer tank/pump 42 a, the quantityof propellant transferred can also be determined based on the knowntransfer tank volume and by simply counting the number of transfercycles. This can be also be used as a back-up means of propellantmetering. Performing the first half cycle filling of the transfer tankcan be used periodically (i.e., at the beginning of direct transferoperations for refueling each client) to check the performance of theflow meter.

Once the desired quantity of propellant is transferred by either of theabove transfer methods, then the transfer tank pressurant can be ventedto space (by opening IVG₃) and then the flex hose pressure relieved (byopening IVL₂) in order to facilitate disconnection of the refueling toolfrom the client satellite FDV.

It is noted that the two (2) methods of transferring fluids from theserver spacecraft storage tank to the client satellite being refueledhave been described with reference to the monopropellant subsystem 16.However, it will be appreciated that the same two methods will beemployed, depending on the circumstances, for the propellant transfersubsystems 12 and 14 for transferring the bipropellant.

Given the financial value of transporting the propellant into space, itis highly preferred that the amount of propellant being provided to eachclient spacecraft is measured accurately to demonstrate that thecustomer has received the full amount of propellant being purchased. Toensure this, two features are incorporated into the present design,first, the system is configured to provide highly regulated flow ofpropellant from the servicer spacecraft to the client satellite which isfacilitated by flow meters installed in the flow paths through thepiping system to the client storage tank.

Secondly, leak detection throughout the flow path is enabled through themonitoring of pressure changes over periods of time where the pressurewould only change due to the presence of a leak (i.e., leak detectionperiod before and after refueling). These monitoring periods may includeperiods of time both before and after opening the backup fill/drainvalve 52 once it has been mated to the client satellite fill/drain valve402, before and after opening the client fill/drain valve, and beforeand after closing the client satellite fill/drain valve aftertransferring propellant to the client satellite, and after closing thebackup fill/drain valve following closure of the client fill/drainvalve.

Each of the valves of the propellant transfer system that must be openedor closed prior to, during, and after, propellant transfer (e.g. FVV,HPLV, HPT, IVG, IVL, NTD, and PR) are actuated by drive circuits thatare interfaced with the above-mentioned command and control system whichis discussed in more detail hereinafter.

A feature of the present invention is the flow control systemscapability to detect the pressure in client propellant tanks 24 and 26once the client fill/drain valve 402 has been opened. Thus, if theclient satellite propellant tank 26 is at a pressure greater than thatof the servicer propellant storage tank up to the as the set point forthe pressure regulators in the high pressure section, then the computeris programmed to use the method illustrated in FIGS. 7 a and 7 b toprovide regulated gas and propellant flow rates. On the other hand, ifthe client satellite propellant tank 26 is at a pressure below that ofthe servicer spacecraft's propellant storage tank 36 then the computeris programmed to use the method illustrated in FIG. 6 using the back-upFDV to regulate the flow rate as appropriate.

As noted above, since the transfer tanks 40 a and 42 a, in combinationwith their associated valves, are essentially a pneumatically poweredpositive displacement pumps, and may be replaced by some other form ofpump such as a motor-driven piston pump, gear pump, or diaphragm pump.

Thus, the present system is designed to have the ability to dynamicallyregulate the flow rates of gas and propellant based on feedback from theflow meter, giving the present system a significant advantage andobviating the need to physically modify the client satellite.

The present propellant transfer system can be used to transferpropellant anywhere, say to a second depot in orbit or on the moon, aswell as satellites needing to be refueled in addition to the systembeing used to remove propellant from dead satellites as mentionedpreviously.

FIGS. 13 and 13 a shows an alternative embodiment of a fluid transfersystem using a mechanical pump. FIG. 13 shows the basic concept of usingan inline pump which in an embodiment is realized by modifying thesystem of FIG. 6 to replace some of the “refill and transfer” plumbingand replace it with inline pump 400. FIG. 13 a shows an embodiment usingredundant pumps 400 located in the parallel flow paths with isolationvalves so that only one pump 400 is active at one time. Only oneisolation valve required to isolate a redundant string, but the Figureshows both entrance/exits of the pumps isolated so that fluid can beexcluded from the redundant pump cavities as needed.

It is noted that all embodiments of the fluid transfer system includethe pressurized gas tank with the inert gas. For the embodiments of thesystem without the inline pump, the pressurized as fulfills two roles,the first to provide the motive force for moving the propellants fromthe servicer satellite to the client satellite, it also and secondly itreplaces the volume that is voided in the servicer satellite storagetank(s). In the case of the embodiments in FIGS. 13 and 13 a that uses amechanical pump as the motive force, the pressurized gas is still usedto replace the volume that is voided from the servicer storage tank(s).Thus while both embodiments use the pressurized gas, the system of FIGS.13 and 13 a uses a lesser amount.

The propellant transfer system disclosed herein may form part of asatellite refueling system which may include a dedicated servicerspacecraft on which the propellant transfer apparatus, including a toolcaddy, robotic arm and various tools, are mounted. FIG. 8 is a blockdiagram showing those items pertaining to the refueling of a satellitein addition to the refueling system. These include a host servicerspacecraft 400, the client satellite 401 to be refueled, the clientvalve(s) 402, a robotic arm 403, the refueling tool 50 releasiblygripped by the end effector of robotic arm 403, a propellant couplingmechanism 405, the propellant outlet hose 32, the propellant transfersystem 10, and a communication system 410 to provide a two-way radiolink 407 to Earth 408. It also shows the stowage points for the backupfill/drain valve 52, the replacement seal fitting 133, the secondaryseal fitting 122 and the reset post 111.

Such a dedicated servicer spacecraft may include a spacecraft dockingmechanism such as that disclosed in U.S. Pat. No. 6,969,030 issued Nov.29, 2005, which patent is incorporated herein in its entirety byreference.

The satellite refueling system includes a multifunction tool asdisclosed in co-pending U.S. patent application Ser. No. 13/652,339filed Oct. 15, 2012, to Roberts et al. (United States Patent Publication2013/XXX) (which is incorporated herein in its entirety by reference)the purpose of which is to provide tool tips needed to gain access tothe fill/drain valves 402 themselves and includes a tool holder and asuite of tool tips which are held by the tool holder and activated by asingle motive source under robotic control. The multifunction tool isreleasibly graspable by the end effector of the robotic arm 403.

The refueling tool 50 for accessing the fill/drain valves 402 on theclient satellite to allow for the transfer of bi- or mono-propellantsfrom the servicer spacecraft to the client satellite may be the same asthat disclosed in co-pending U.S. patent application Ser. No.61/566,893, filed as a full utility application Serial No. XXX filedDec. 5, 2012 (United States Patent Publication XXX) (which isincorporated herein in its entirety by reference).

The command and control system is configured to control movement of therobotic arm 403 and the end effector attached thereto for controllingthe action of the multifunction tool, as well as the refueling tool 50.This may be the same command and control system mentioned above that isinterfaced with the flow control system, for example a computer mountedon the servicer satellite which is programmed with instructions to carryout all operations needed to be performed by the servicer satelliteduring approach, capture/docking with the client satellite and refuelingoperations. It may also be a separate computer system. The satelliterefueling system includes a vision system for viewing the operation ofthe multifunction tool and the refueling tool during propellant transferoperations. Communication system 410 is interfaced with the robotic arm403 and configured to allow remote operation (from the Earth 408 or fromany other suitable location) of the vision system (which may include oneor more cameras), the robotic arm 403 and hence the tools. The visionsystem may include distinct markers mounted on the fluid transfercoupling used to couple the fluid transfer system storage tank andpiping system to the fill/drain valve of the client satellite, as wellas markings on all tools associated with the fluid transfer operation.

These cameras may be used within a telerobotic control mode where anoperator controlling the servicing robotics on earth views distinctviews of the worksite on display screens at the command and controlconsole. In an alternative mode, the position of elements like the filldrain valve may be determined by either a stereo camera and visionsystem which extracts 3D points and determines position and orientationof the fill-drain valve or other relevant features on the worksite fromwhich the robotic arm holding tools (multi-function tool, refuelingtool) can be driven to these locations according the sensed 6degree-of-freedom coordinates.

The stereo camera could also be replaced with a scanning or flash lidarsystem from which desired 6 degree-of-freedom coordinates could beobtained by taking measured 3-D point clouds and estimating the pose ofdesired objects based on stored CAD models of the desired features orshapes on the refueling worksite. For those applications where thespacecraft was designed with the intention to be serviced, a simpletarget such as described in Ogilvie et al. (Ogilvie, A., Justin Allport,Michael Hannah, John Lymer, “Autonomous Satellite Servicing Using theOrbital Express Demonstration Manipulator System,” Proc. of the 9thInternational Symposium on Artificial Intelligence, Robotics andAutomation in Space (i-SAIRAS '08), Los Angeles, Calif., Feb. 25-29,2008) could be used in combination with a monocular camera on theservicing robotics to locations items of interest such as the fill-drainvalve. Finally, the robotic arm or device used to position the devicemay include a sensor or sensors capable of measuring reaction forcesbetween the tools and the work-site (e.g. fill-drain valves). These canbe displayed to the operator to aid the operator in tele-operationcontrol or can be used in an automatic force-moment accommodationcontrol mode, which either aids a tele-operator or can be used in asupervised autonomous control mode.

A system of this type is very advantageous particularly for space-basedsystems needing remote control. The various components making up therefueling system may be retrofitted onto any suitable satellite to beused as a servicer spacecraft for 400 refueling. The servicer spacecraftwith the propellant transfer apparatus mounted thereon could be carriedon a larger “mother ship” and launched from there or stored on anorbiting space station and launched from there when needed.

The system may be operated under tele-operation by a remotely locatedoperator, for example located on earth, in the “mother ship”, or in anorbiting space station. Under pure remote teleoperator control, commandand control is by the teleoperator who may issue commands directly tothe propellant transfer system to open and close selected fluid and gasvalves based on the known client pressure and feedback from the flowmeter.

In addition, the servicer satellite 400 (FIG. 8) includes an onboardcomputer control system 500 which may be interfaced with the propellantflow control system, shown at 560 in FIG. 12 so that it can drive allthe components that are opened and closed during the propellant transferoperations in a selected sequence depending on which mode of propellanttransfer has been selected based on the pressure in client tank 26. Withthe presence of the computer control system 500 interfaced with thepropellant flow control system, the propellant transfer process may beautonomously controlled by a local Mission Manager or may include somelevels of supervised autonomy so that in addition to being under pureteleoperation there may be mixed teleoperation/supervised autonomy.

The present system is also configured for full autonomous operation. Afully autonomous system is a system that measures and responds to itsexternal environment; full autonomy is often pursued under conditionsthat require very responsive changes in system state to externalconditions or for conditions that require rapid decision making forcontrolling hazardous situations. The implementation of full autonomy isoften costly and is often unable to handle unforeseen or highlyuncertain environments. Supervised autonomy, with human operators ableto initiate autonomous states in a system, provides the benefits of aresponsive autonomous local controller, with the flexibility provided byhuman tele-operators.

The method of resupplying propellant to a client satellite comprises thefollowing steps. First, servicer spacecraft 400 is maneuvered intoposition and orientation with respect to client satellite 401 under a)computer control, based on sensor feedback from onboard sensors havingoutputs interfaced with the onboard computer system on the servicerspacecraft, or under b) teleoperator control by an operator remotelylocated from the servicer spacecraft, or c) under combined computer andteleoperator control.

Once the servicer spacecraft 400 and the client satellite 401 are in thedesired relative position and orientation, servicer spacecraft 400releasibly captures client satellite 401. The capture process mayinvolve direct capture by for example and satellite capture device suchas that disclosed in U.S. Pat. No. 6,969,030 issued Nov. 29, 2005, whichpatent is incorporated herein in its entirety by reference.Alternatively, the client satellite 401 may be realisibly captured by adedicated robotic arm 430 (shown in dashed outline in FIG. 8). Or, acombination capture process may include use of both a robotic arm 430and the mechanism shown U.S. Pat. No. 6,969,030.

After capture is complete, robotic arm 403 is deployed to releasiblygrasp a multifunction tool 440, such as disclosed in co-pending U.S.patent application Ser. No. 61/546,770 filed Oct. 12, 2011, to Robertset al. (United States Patent Publication 2013/XXX), whereupon roboticarm 403 is commanded to use tool 440 (see FIG. 12) to gain access to thepropellant tank valves 402 on client satellite 401 by cutting throughany thermal blankets and lock wires on the propellant valve(s) 402. Oncethis is complete, multifunction tool 440 is released and sequesteredonto its tool holder by the end effector on the servicer spacecraft 400and the refueling tool 50 is releasibly gripped by the end effector ofrobotic arm 403. The refueling tool 50 is then commanded to releasiblygrasp backup fill/drain valve 52 from its holder on satellite 400 and ismated with client satellite propellant tank fill/drain valve 402.

FIG. 9 shows the client satellite drain/fill valve 402, the secondaryseal fitting 117, the valve actuation nut 131, the torque reactionfeatures 121 and the client valve centreline 140. FIG. 10A shows thebackup fill/drain valve 52 with the secondary seal threaded feature 127,the backup fill/drain valve actuation nut 132, the backup fill/drainvalve torque reaction feature 129 and the backup fill/drain valve sealfitting 128.

FIG. 10B shows the replacement secondary seal fitting 134 with the sealfitting threaded feature 137, the replacement secondary seal fittingtorque reaction feature 135 and the seal fitting 136. FIG. 11 shows thebackup fill/drain valve 52 engaged with the client satellite drain/fillvalve 402 once it has been coupled thereto by the refueling tool 50.

After the client satellite propellant tank has been filled with thedesired amount of propellant, if the client satellite fill/drain valve402 is successfully closed (i.e., does not leak) then backup fill/drainvalve 52 is removed by refueling tool 50 and refueling tool 50 grasps asecond seal fitting 134 which is then mated to the fill/drain valve 402,which is then left behind. If, upon testing for leaks in the systemafter the fill/drain valve 402 is closed, valve 402 is found to leak,then the backup fill/drain valve 52 is closed, and then left behindmated to valve 402 and the secondary seal fitting 134 is attached to thebackup fill/drain valve 52 by tool 50 and it too is left behind.

The thermistors are used to measure the temperature of associatedequipment on all the propellant transfer subsystems and hardwaresupporting the subsystems. Thermistors may be included in the refuelingsubsystem as follows.

-   -   1) For all tanks, the thermistors are used to monitor the        temperature. These temperature sensors/thermistors, along with        the high pressure transducers, are used to check that the tanks        are in a safe state. The thermistors and high pressure        transducers, being interfaced with the computer control system        (discussed below), are also used to calculate the pressure        difference between interconnected tanks to predict the flow rate        (and direction of flow) between them.    -   2) For the pressurant tanks, the thermistors, along with the        high pressure transducers these are used to determine the        quantity of pressurant remaining using P·V=m·R·T.    -   3) For the transfer tanks, they are used to determine the        propellant density and the tank volume as a function of        temperature. The propellant mass transferred is the given by:        (number_of_cycles)*(TT_volume)*(propellant_density).    -   4) For the pressure transducers the thermistors are used to        correct the pressure measurement as a function of temperature.    -   5) For the flow meter transducers, the thermistors are used to        correct the flow measurement as a function of temperature. The        thermistors are also used to determine the propellant density as        a function of temperature. The flow meter measures the volume        flow rate which is then multiplied by the propellant density to        determine the mass flow rate. The computer control system,        discussed below, is interfaced with all the sensors and is        programmed with instructions to make the above calculations and        to monitor all parts of the system.

The use of backup fill/drain valve 52 is very advantageous for severalreasons. If the client fill drain valve cannot be properly closed (i.e.,if it leaks), then propellant would leak from the client propulsionsubsystem as soon as the refueling tool is disconnected. This situationis not acceptable and the risk of its occurrence could dissuadepotential customers from consider propellant resupply of theirspacecraft. This risk is mitigated through the use of the backupfill/drain valve 52. The system may be configured in such a way thatbackup fill/drain valve 52 also functions as a throttling valve. If thebackup fill/drain valve 52 is only partially opened, then it hasincreased flow resistance and so can be used to regulate the flow rateof propellant. This is useful for the direct transfer method asotherwise the flow rate would be determined solely by the pressuredifference between the client propellant tank and the servicerspacecraft propellant storage tank. The backup fill/drain valve 52provides a secondary sealing of the principal flow path from the clientstorage tank 26 to the vacuum of space.

Referring now to FIGS. 8 and 12, an example computing system 500 formingpart of the propellant resupply system is illustrated. The systemincludes a computer control system 525 configured, and programmed tocontrol movement of the robotic arm 403 during the entire procedure ofaccessing the client satellite fill/drain valve 402, attachment of thebackup fill/drain valve 52, mating of propellant outlet hoses 32 or 34to the backup fill/drain valve 52, transfer of propellant into tanks 22or 26, demating hoses 32 or 34 from backup fill/drain valve 52, sealingvalve 52 and decoupling of the servicer spacecraft 400 from the clientsatellite 401.

As mentioned above, computer control system 525 is interfaced withvision system 550, the flow control system 560 of the propellanttransfer system, and robotic arm 403. Previously mentioned communicationsystem 410 is provided which is interfaced with the robotic arm 403 andconfigured to allow remote operation (from the Earth 408 or from anyother suitable location) of the vision system (which may include one ormore cameras 550), the robotic arm 403 and the flow control system 560.A system of this type is very advantageous particularly for space basedsystems needing remote control. By providing a suite of tool tips in atool caddy that are accessible to the multifunction tool 540 (FIG. 12)that are configured to be activated by a single motive source on themultifunction tool (not shown) such that they do not need their ownpower sources provides an enormous saving in weight which is a premiumon every launch.

Some aspects of the present disclosure can be embodied, at least inpart, in software. That is, the techniques can be carried out in acomputer system or other data processing system in response to itsprocessor, such as a microprocessor, executing sequences of instructionscontained in a memory, such as ROM, volatile RAM, non-volatile memory,cache, magnetic and optical disks, or a remote storage device. Further,the instructions can be downloaded into a computing device over a datanetwork in a form of compiled and linked version. Alternatively, thelogic to perform the processes as discussed above could be implementedin additional computer and/or machine readable media, such as discretehardware components as large-scale integrated circuits (LSI's),application-specific integrated circuits (ASIC's), or firmware such aselectrically erasable programmable read-only memory (EEPROM's).

FIG. 12 provides an exemplary, non-limiting implementation of computercontrol system 525, forming part of the command and control system,which includes one or more processors 530 (for example, aCPU/microprocessor), bus 502, memory 535, which may include randomaccess memory (RAM) and/or read only memory (ROM), one or more internalstorage devices 540 (e.g. a hard disk drive, compact disk drive orinternal flash memory), a power supply 545, one more communicationsinterfaces 410, and various input/output devices and/or interfaces 555.

Although only one of each component is illustrated in FIG. 12, anynumber of each component can be included computer control system 525.For example, a computer typically contains a number of different datastorage media. Furthermore, although bus 502 is depicted as a singleconnection between all of the components, it will be appreciated thatthe bus 502 may represent one or more circuits, devices or communicationchannels which link two or more of the components. For example, inpersonal computers, bus 502 often includes or is a motherboard.

In one embodiment, computer control system 525 may be, or include, ageneral purpose computer or any other hardware equivalents configuredfor operation in space. Computer control system 525 may also beimplemented as one or more physical devices that are coupled toprocessor 530 through one of more communications channels or interfaces.For example, computer control system 525 can be implemented usingapplication specific integrated circuits (ASIC). Alternatively, computercontrol system 525 can be implemented as a combination of hardware andsoftware, where the software is loaded into the processor from thememory or over a network connection.

Computer control system 525 may be programmed with a set of instructionswhich when executed in the processor causes the system to perform one ormore methods described in the present disclosure. Computer controlsystem 525 may include many more or less components than those shown.

While some embodiments have been described in the context of fullyfunctioning computers and computer systems, those skilled in the artwill appreciate that various embodiments are capable of beingdistributed as a program product in a variety of forms and are capableof being applied regardless of the particular type of machine orcomputer readable media used to actually effect the distribution.

A computer readable medium can be used to store software and data whichwhen executed by a data processing system causes the system to performvarious methods. The executable software and data can be stored invarious places including for example ROM, volatile RAM, non-volatilememory and/or cache. Portions of this software and/or data can be storedin any one of these storage devices. In general, a machine readablemedium includes any mechanism that provides (i.e., stores and/ortransmits) information in a form accessible by a machine (e.g., acomputer, network device, personal digital assistant, manufacturingtool, any device with a set of one or more processors, etc.).

Examples of computer-readable media include but are not limited torecordable and non-recordable type media such as volatile andnon-volatile memory devices, read only memory (ROM), random accessmemory (RAM), flash memory devices, floppy and other removable disks,magnetic disk storage media, optical storage media (e.g., compact discs(CDs), digital versatile disks (DVDs), etc.), among others. Theinstructions can be embodied in digital and analog communication linksfor electrical, optical, acoustical or other forms of propagatedsignals, such as carrier waves, infrared signals, digital signals, andthe like.

The present system is advantageous over the system disclosed in ScottRotenberger, David SooHoo, Gabriel Abraham, Orbital Express FluidTransfer Demonstration System in Sensors and Systems for SpaceApplications II, edited by Richard T. Howard, Pejmun Motaghedi, Proc. ofSPIE Vol. 6958, 695808, (2008) for the following reasons. Rotenbergerneeds to fill and pressurize a transfer tank to 50 Pounds per SquareInch Differential. (psid) and perform unregulated blow down to thedesired client fill and pressure by executing a number of “blow andrefill” cycles. Through this unregulated approach, there is asignificant difference between the initial transfer pressure and finalpressure when the transfer cycle is about to be completed. Consequentlythere is a large difference between initial flow rate and the final flowrate (i.e., large dynamic range), making accurate mass flow ratemeasurement difficult to achieve.

The present systems maintains a virtual storage tank preferably at about20 bar (but is not restricted to this pressure) and directly regulatesthe exit flow of propellant so that the transfer flow rate is safe andthe pressure is increased in the client satellite in a continuous,controlled, and slow manner. The fluid transfer system disclosed hereinconserves pressurant in this way and it is possible to regulate the flowto be whatever is required at any time during the propellant transferprocedure. A flow meter measures the flow rate and this can be useddirectly for feedback.

In addition, the flow rate regulation features incorporated abovepreclude the need to modify features of the client propulsion systemsuch changing the client tank inlet configuration as Rotenberger et al.were required to do.

While the system disclosed herein has been described as a remotepropellant transfer system for resupplying satellites in orbit once thepropellant has been depleted or partially depleted, it will beunderstood the present system may also be used for safely transferringpropellant to satellites prior to being launched into space. In thiscontext, no direct human contact is needed during the fueling proceduresince the whole process can be remotely controlled from a safe distance.In this situation the fueling is done prior to sealing the thermalblankets and wiring the valves with the various lockwires so these wouldnot have to be cut off in the pre-launch scenario.

The specific embodiments described above have been shown by way ofexample, and it should be understood that these embodiments may besusceptible to various modifications and alternative forms. It should befurther understood that the claims are not intended to be limited to theparticular forms disclosed, but rather to cover all modifications,equivalents, and alternatives falling within the spirit and scope ofthis disclosure.

Therefore what is claimed is:
 1. A method of refueling a clientsatellite, comprising: a) maneuvering a servicer spacecraft into closeproximity to a client satellite in a position and orientation suitableto releasibly couple the client satellite to the servicer spacecraft,releasibly coupling the client satellite to the servicer spacecraft; b)deploying, and commanding, a robotic arm mounted on the servicerspacecraft to releasibly grasp a multifunction tool, commanding therobotic arm to manipulate the multifunction tool to access a fill/drainvalve on the client satellite in flow communication with a storage tankon the client satellite; c) commanding the robotic arm to sequester themultifunction tool and releasibly grasp a refueling tool, commanding therobotic arm to manipulate the refueling tool to releasibly grasp apropellant outlet hose connected to a propellant transfer system andmate it to the fill/drain valve; e) opening the fill/drain valve andmeasuring a pressure in the client satellite storage tank and based onthe measured pressure, configuring a flow control system mounted on theservicer spacecraft to transfer propellant under regulated pressure andflow rate conditions suitable for the measured pressure; f) commandingthe configured flow control system of the propellant transfer system totransfer propellant from a propellant storage tank on the servicerspacecraft through a piping system to the propellant hose to the storagetank on the client satellite; g) once a desired quantity of propellanthas been transferred to the client satellite, commanding the robotic armto manipulate the refueling tool to close the fill/drain valve; h)demate the propellant hose from the fill/drain valve and sequester it;and i) sequestering the robotic arm and decoupling the servicerspacecraft from the client satellite.
 2. The method according to claim 1wherein in step c) prior to commanding the robotic arm to mate thepropellant hose to the fill/drain valve, commanding the robotic arm tomanipulate the refueling tool to releasibly grasp a backup fill/drainvalve and mate the backup fill/drain valve with the fill/drain valve onthe servicer spacecraft, and thereafter commanding the robotic arm tomanipulate the refueling tool to releasibly grasp the propellant hoseand mate it to the backup fill/drain valve and to open both thefill/drain valve and the backup fill/drain valve.
 3. The methodaccording to claim 2 including monitoring pressure over periods of timein the piping system after the fill/drain valve and the backupfill/drain valve are opened for detection of leaks.
 4. The methodaccording to claim 3 wherein the periods of time include periods of timeboth before and after transferring propellant to the client satellite.5. The method according to any one of claim 1 wherein in step e) ofmeasuring the pressure of the client satellite storage tank, if thepressure is lower than that of the servicer spacecraft's propellantstorage tank, the step of configuring a flow control system mounted onthe servicer spacecraft to dispense propellant under regulated pressureand flow rate conditions suitable for the measured pressure includesproviding a first flow path through the propellant transfer system andproviding regulated flow of i) pressurized gas between at least onepressurized gas tank and at least one propellant storage tank containedin the propellant transfer system, and ii) propellant between said atleast one propellant storage tank and said propellant outlet hose. 6.The method according to claim 5 wherein in step e) of measuring thepressure of the client satellite storage tank, if the pressure is higherthan that of the servicer spacecraft's propellant storage tank up to aset point for pressure regulators in a high pressure section pipingsystem associated with said at least one pressurized gas tank, the stepof configuring a flow control system mounted on the servicer spacecraftto dispense propellant under regulated pressure and flow rate conditionssuitable for the measured pressure includes providing a second flow paththrough the propellant transfer system and providing regulated flow ofpropellant from at least one storage tank to a fluid volume of at leastone propellant transfer tank facilitated by venting a gas volume of saidat least one transfer tank to space and opening a flow path between atleast one propellant storage tank and said fluid volume, and after adesired quantity of propellant has transferred to said fluid volume thencease venting said gas volume to space, and thereafter pressurizing thegas volume and opening a flow path from said fluid volume to saidpropellant outlet hose thereby forcing the propellant to flow from saidfluid volume to said propellant outlet and into the client satellitestorage tank.
 7. The method according to claim 2 wherein step g) furtherincludes the steps of commanding the robotic arm to manipulate therefueling tool to demate the backup fill/drain valve from the clientsatellite fill/drain valve and sequester it, and commanding the roboticarm to manipulate the refueling tool to releasibly grasp a replacementsecondary seal fitting and couple it to the client satellite fill/drainvalve.
 8. The method according to claim 1 wherein step g) furtherincludes the steps of commanding the robotic arm to manipulate therefueling tool to releasibly grasp a replacement secondary seal fittingand couple it to the backup fill/drain valve thereby leaving behind thebackup fill/drain valve still mated to the client satellite fill/drainvalve.
 9. The method according to claim 1 wherein said client satelliteuses a monopropellant, and wherein monopropellant is transferred fromsaid servicer spacecraft to said client satellite.
 10. The methodaccording to claim 1 wherein said client satellite uses a bipropellantincluding a fuel and an oxidizer, and wherein steps b), c), d), e), f),g) and h) are first carried out for transferring fuel to an associatedfuel propellant tank on the client satellite, and then steps b), c), d),e), f), g) and h) are repeated for transferring oxidizer to anassociated oxidizer propellant tank on the client satellite.
 11. Themethod according to claim 1 wherein said commands are issued through acommunication system configured to provide communication between saidcommand and control system and a remote operator for remoteteleoperation control, supervised autonomous control, or fullyautonomous control of propellant transfer operations between theservicer spacecraft and the client satellite.